Means for stabilizing fluid flow in diffuser-combustor systems in axial flow gas turbine engines



Jan. 23, 1968 J s. ALFORD 3,364,678

MEANS FOR STABILIZING FLUID FLOW IN DIFFUSER-COMBUSTOR SYSTEMS IN AXIALFLOW GAS TURBINE ENGINES Filed Feb. 28, 1966 '1 EH m w [ff 5 2 I v. IINVENTOR.

United States Patent 3,364,678 MEANS FOR STABILIZING FLUID FLOW IN DIF-FUSER-COMBUSTOR SYSTEMS IN AXIAL FLOW GAS TURBENE ENGINES Joseph S.Alford, Wyoming, Ohio, assignor to General Electric Company, acorporation of N ew York Filed Feb. 28, 1966, Ser. No. 530,599 8 Claims.(Ci. 6039.65)

ABSTRACT OF THE DISCLGSURE The disclosure shows a coupling between adiffuser and the inlet to a combustor wherein the combustor comprisescasing members and liners providing for the flow of cooling air aroundthe combustor as well as entry of primary combustion air into thecombustor. The provision of an abrupt change in angle between thediffusers and the casings and specified relationships between the casins and liners provides a controlled anchor point for air flow as itseparates in passing from the diffuser to the spacing between thecasings and liners. This results in greater stability of operation.

Gas turbine engines of the type referred to comprise an axial flowcompressor which provides pressurized air to a continuous fiowcombustor. The present invention deals particularly with enginesemploying annular combustors comprising inner and outer walls and a pairof spaced liners therebetween which define three concentric flow paths.Normally air discharged from the compressor passes through a divergent,annular passageway which is known as a diffuser. From the diffuser theair stream is split into the three flow paths through the combustor.Combustion of fuel is maintained in the central flow path between theliners, while the outer flow paths provide cooling air and alsoadditional air for supporting further combustion of the fuel downstreamof the liner.

In many compressors circumferential pressure gradients exist in thepressuried air stream discharged from the compressor. Pressure gradientsmay also be attributed to radial support struts commonly extendingthrough the diffuser. Such pressure gradients tend to fluctuate duringoperation of the engine and cause changes in the points where the airstream attaches to the diffuser walls. Such fluctuation causesself-excited accoustical oscillation and vibration which are harmful tothe engine structure and to engine operation. These fluctuations alsocause or accentuate circumferential, thermal gradients and result in hotstreaks which can greatly reduce the life of the combustor and/ or theengine turbine.

The object of the invention is to minimize, if not eliminate, the airflow instablility problems in gas turbine engines discussed above.

The above, brief description of air flow from a diffuser to a combustorin a gas turbine engine sets the environment for the present invention.This environment may be more specifically described by pointing out thatthe combustor walls and adjacent liners are respectively mutuallyconvergent to define entrances to the outer fiow paths which have anominal divergent direction from the central flow path through thecombustor.

In its broader aspects the invention is characterized by the diffuserhaving a ratio of outlet to inlet area below a value required to insurethat lack of any fiow separation occurs along the diffuser walls. Theconvergent portion of at least one combustor wall joins the diffuserwall at a sharp angle or break and the convergent portion of theadjacent liner defines, in combination therewith the air flow pathhaving a nominal divergent direction. The specified adjacent linerextends inwardly at least to a plane normal to the nominal divergentflow path and in- 3 ,364,578 Patented Jan. 23, 1968 tersecting the sharpangle of the juncture of the diffuser and combustor walls.

In this fashion the air flow is anchored to the diffuser Wallimmediately upstream of the sharp break where the combustor wall joinsit. immediately downstream and along the divergent flow path from thissharp break, a controlled separation occurs. Thereafter as the divergentflow reattaches itself to the combustor wall and a stabilized conditionis obtained wherein the controlled separation which is maintainedcircumferentially, minimizes, if not eliminates, vibrations and thermalhot streaks resulting from pressure gradients primarily attributable tothe characteristics of the axial fiow compressor supplying pressurizedair for the combustor.

These and other objects, features and advantages of my invention will beapparent from the following detailed description taken together with theaccompanying drawings in which:

FIG. 1 is a schematic view in cross-section of a typical axial-flow gasturbine engine in which the invention has been incorporated;

FIG. 2 is a partial view, enlarged and in cross-section, of thecompressor discharge area, diffuser and general inlet area of thecombustor chamber of the engine of FIG. 1; and

FIG. 3 is a further enlarged view of the junction between the dilfuserand combustor illustrating the air fiow therebetween.

Turning now specifically to FIG. 1 indicated generally at 10 is a gasturbine engine of the well-known turbojet variety. While this is used asan example to illustrate the invention it will be understood that theinvention has application to any apparatus utilizing a continuous fluidflow combustion system, for example, aircraft turbofan or land-basedcombustion engines. in any event, as shown in FIG. 1 the turbojet it)includes an outer housing 11 having an inlet end 12, receiving air whichenters a multistage axial flow compressor 14 having rows of rotor blades16. Interspersed with the rotor blade rows are rows of stator blades 18.The stator blades are affixed at one end to the inner surface of thehousing 11. At the downstream end of the compressor is a row ofcompressor outlet guide vanes (OGVs) 20, followed by an annular diffuserpassage or compressor discharge passage indicated generally at 22.

The compressor discharge passage or diffuser duct 22 comprises a pair ofconcentric inner and outer walls 24 and 26, respectively, divergent in adownstream direction. Struts 27 (FIG. 2) span the diffuser and comprisepart of the structural components of the engine in normal fashion. Thediffuser discharges the pressurized air into a combustor indicatedgenerally at 30 from whence the heated gases exit at high velocitythrough the power turbine 32. The power turbine extracts work to drivethe compressor 14 by means of a connecting shaft 34 on which bothcomponents are mounted. The rotating compressor-turbine set is mountedin the engine by suitable bearing means. The hot gas stream leaving theturbine is discharged to atmosphereto provide thrustthrough an exhaustnozzle which may be of the adjustable type, as shown at 33 in FIG. 1.

Referring now to the enlarged view of FIG. 2, it will be seen that thecombustor comprises an outer casing wall, indicated generally at 48',and an inner casing wall,

' indicated generally at 42. The outer and inner walls are spaced from apair of outer and inner combustion chamber liners, indicated generallyat 44 and 46 respectively. The liner Wall members are appropriatelysupported in the combustor and are interconnected at their forward endsby an annular wall 48. Intermediate the radially spaced ends of wall 48is an opening 50 adapted to receive a fuel nozzle 52, illustrated inphantom in FIG. 2 and shown in solid lines in PEG. 1. Nozzle 52 issupplied, in the usual manner, with fuel to support combustion. Thecombustor 30, as will be perceived from the drawings is of the annulartype, as contrasted to the can or cannular variety.

It will be apparent that the combustor walls 46, 42 and liners 4 46define three concentric annular flow paths into which air from thediffuser is split. These flow paths and the structure defining them willbe described simply as to their relative relations shown in thelongitudinal half section and fragmentary half section found in FIGS. 2and 3.

It will be noted that the forward ends of the liners 4- have cowlportions 44a, 45a respectively, terminating in spaced lips 441), 4-6!)which define the entrance for a central portion of the diffuser airstream into the open ended primary combustion chamber defined by theseliners. The lips 44]), 461) are preferably rounded and somewhatthickened for strength purposes. The forward ends of the combustor Wallsalso have mutually convergent portions 46a, 42a and are joined to theends of the diffuser walls 26, 24 at 49b, 42!; preferably in a commonplane (X) normal to the axis of the engine.

The combustor wall 40 and liner 44 and also liner 46 and combustor wall42 define flow paths 6%, 62 for the diffuser air which are firstdivergent from the diffuser flow path and then curved to thelongitudinally extending downstream portions of the liners. The outerflow paths 60, 62 provide cooling air for the combustor walls as well assupplying additional combustion air through liner openings 64 downstreamof the nozzle 52, as seen in FIG. 1.

The present invention involves the manner in which air from the diffuseris diverted or turned into the outer flow paths 50, 62. The means for sodiverting air into flow path 60 will be described with reference to FIG.3 with the understanding that such means are correspondingly found inflow path 62.

Generally speaking turning of the air stream is characterized by theprovision of a sharp break in the outer walls where the turn is made,i.e., the juncture of diffuser wall 24- with end 4% of the combustorwall. Further it is preferred that the angle a between the diffuser wall24 and convergent combustor wall portion be approximately between 135and 145.

The entrance portion to flow path 6% is defined also by liner cowlportion 44a which, in a plane (B) extending from edge 4%, normal to theinitial, nominal divergent flow direction (arrow V) is spaced from wallportion 49a a distance between .8 and 1.0 h where k equals half of theheight H of the diifuser outlet passage in plane X. It is alsopreferable that the cowl lip 44b extend into the diffuser air flow pathbeyond plane B a further distance of between .1 and .3 it.

Another preferred relation is that the wall portions 40a and 44a beparallel and fiirther remain straight for a distance of approximately 2h. Thereafter these walls are gently curved to the illustrateddownstream portions.

With the described relationships the air stream is diverted into flowpath on in a fashion which at all times maintains air separation at afixed point, thereby eliminating, or at least greatly reducing, anyundesirable shifting which can cause self excited acoustical vibrationand other harmful effects.

As previously indicated the diffuser is chosen with an outlet to inletarea ratio between 1.6 and 1.7. This relation is made because withinthese limits it has been found that for most axial flow compressorsthere is little or no separation of the air stream from the diffuserwalls 24, 26 and yet a sufiicient pressure increase is obtained foreffective combustor operation, particularly in combination with thedescribed fiow turning means.

Assurance is thus had that the air st earn is anchored to the diffuserwall 2 immediately upstream of the sharp break at the forward edge 40bof the combustor wall.

As the air stream is turned by cowl portion 44a toward the nominaldirection of arrow V the air stream separates from wall portion dtlaimmediately downstream of edge dtlb as indicated by reference characters and the air flow lines in FIG. 3. Once turned, it has been found thatthe diverted air stream will reattach itself to the combustor wall 46aat a distance of 1 to 2 h from the turning point, particularly where asstraight parallel flow path is defined. Hence it is preferred that theportions dtia and 44a have a minimum length of 2 h from plane B.

With the diverted air stream reattached to the walls defining its flowpath, it can then be curved back to a longitudinal direction along thedownstream portions of the combustor. By having the diverted streamanchored and separated at fixed longitudinal points extendingcircumferentially of the annular flow path, pressure gradicuts andresultant vibrations and thermal hot streaks are greatly minimized, ifnot eliminated. These ends are further attained with a minimum of energylosses which is attributed to the specified relationships of the cowlportions 44a and 46a.

The scope of the inventive concepts herein described is not necessarilylimited to the specific structure described but is to be derived fromthe following claims.

Having thus described the invention what is claimed as novel and desiredto be secured by Letters Patent of the United States is:

1. In a gas turbine engine having a diffuser for receiving air from acompressor and a combustor to which air is directed from the diffuser tosupport combustion theresaid diffuser comprising mutually divergentinner and outer walls defining an annular flow path,

said combustor comprising inner and outer walls respectively connectedto said diffuser walls and a pair of spaced liners spaced therebetweenand defining three concentric flow paths,

the forward ends of said liners having mutually convergent cowl portionsterminating in spaced lips defining the entrance to the central one ofsaid concentric flow paths,

the forward end portions of said combustor walls also being mutuallyconvergent and defining in combination with adjacent liners the outer ofsaid concentric flow paths, which are divergent from said central flowpath characterized in that the ratio of the outlet to inlet area of saiddiffuser is below a value required to insure lack of any flow separationfrom the diffuser walls,

said convergent portion of at least one combustor wall joins thediffuser wall at a sharp angle and the convergent portion of theadjacent liner defines in combination with the convergent portion ofsaid one combustor wall an air flow path having a nominal divergentdirection, said adjacent liner extending inwardly at least to a plane(B) extending normal to said nominal divergent flow path and from thesharp angle of the juncture of said diffuser and combustor walls.

2. A combination as in claim 1 wherein the angle (a) between the joinedcombustor and diffuser walls is between approximately -145.

3. A combination as in claim 1 wherein the said adjacent combustor walland liners thereto are spaced apart a distance, in said plane (B),between .8 and unity of half the distance (h) between the diffuser wallsat their discharge ends.

4. A combination as in claim 3 wherein the said liner extends upstreamfrom said plane (B) a distance between .1 and .3 of half the distance(11) between the diffuser walls at their discharge ends.

5. A combination as in claim it wherein the conver ent portion of theother combustor wall joins the other diffuser wall at a sharp angle and5 6 the convergent portion of the other adjacent liner dearerespectively spaced apart a distance, in said plane fines, incombination with the convergent portion of (B), between .8 and unity ofhalf the distance (h) said other combustor Wall, a second air flow pathbetween the diffuser walls at their discharge ends. having a nominaldivergent direction, said other adja- S. A combination as in claim 7wherein cent liner extending inwardly at least to a plane (B) 5 theangle (1x) between the joined combustor and difextending normal to saidother nominal divergent fuser walls is between approximately 135-l45 andflow path and from the sharp angle of the juncture the liner wallsextend inwardly from plane (B) a disof said other diffuser and combustorwalls. tance between .1 and .3 of half the distance (It) 6. Acombination as in claim 5 wherein between the diffuser Walls at theirdischarge ends. the convergent combustor wall and liner portionsdefining said divergent paths are straight and parallel and ReferemesCited extend downstream of said plane (B) a distance UNITED STATESPATENTS approximately equal to the distance (H) between the diffuserWalls at the discharge ends thereof- 322:2"? 1811323 iitanrjjjiijjj:23:33:53

7. A combination as in claim 6 wherein 15 the combustor walls and theliners adjacent thereto JULIUS E. WEST, Primary Examiner.

